(1) Field of the Invention
To control a helicopter, in general via the main rotor or tail rotor, the pilot has to send a command which moves the rotor blades. In case of installed servo systems either for comfort purposes or for necessary force amplification due to high air loads, the pilot command will activate an actuator motion.
There are two classical methods for the actuator power generation, electric or hydraulic. The hydraulic method needs a complex system of reservoirs, pumps, filters, valves, lines, etc. and implies important maintenance efforts. Such a hydraulic system is shown as example in FIG. 1 of the present application with reference to the publication: Aircraft Systems, Ian Moir, Allan Seabridge, AIAA Educational Series, 2008, particularly see FIG. 1.12 on page 19. The required reliability of electro-mechanical actuators is an important issue in aeronautics along with the reduction of redundancy and complexity for the usage of such electro-mechanical actuators.
(2) Description of Related Art
The document EP 0325027 (A1) discloses a mechanical actuator assembly which automatically releases the load controlled by the actuator when the actuator jams or otherwise fails. No hint to a fail free actuator as such can be derived from EP 0325027 (A1).
The document U.S. Pat. No. 5,144,851 (A) is directed to a jam tolerant ball screw actuator driven through a differential via plural power paths therein. A hint to a control system is not referred to in U.S. Pat. No. 5,144,851 (A).
The document U.S. Pat. No. 4,137,784 (A) discloses an electromechanical actuator which acts as an analogue of a fluid operated cylinder. A tubular piston moves along the axis of a fixed outer cylinder. One end of the piston is fitted with a threaded stud to which a clevis or other load attachment device can be attached. The object of U.S. Pat. No. 4,137,784 (A) is to provide reliable turn off functionality of the actuator.
The document U.S. Pat. No. 7,100,870 discloses a jam tolerant electromechanical actuating system in an aircraft and a method for controlling this system including: Locating a physical coupling/decoupling mechanism between the load and an actuator assembly as close as practicable to the load; constructing the coupling/uncoupling mechanism to be reversible, and hence testable; and controlling the connection/disconnection via decision making electronics which will detect any system failure by monitoring. The jam tolerant electromechanical actuating system of document U.S. Pat. No. 7,100,870 relies on at least two electromechanical actuator assemblies each of said actuator assemblies being provided with a motor, load and position sensors and a coupling/decoupling mechanism and a disconnect actuator rendering said actuator system heavy and complex.
In conventional usage of hydraulic actuators basically two methods remain for sending the pilot's command to the actuator, either by “Fly By Wire” (FBW), or mechanically by lever and bell crank linkage.
The document US 2004251061 (A1) discloses an active steering system being manually commanded by a mechanical linkage with variable assist including a differential actuator having an input gear and an output gear. The differential actuator has a default kinematic relationship between the input gear and the output gear such that magnitude of an output speed and an output torque is approximately equal to a magnitude of an input speed and an input torque with opposing directions. The system also includes an input device that is in operable communication with the differential actuator and an output device that is in operable communication with the differential actuator. The differential actuator is operable to vary a ratio between the input device and the output device. The system also includes a steering mechanism that is in operable communication with the differential actuator and the steering mechanism is configured to reverse the opposing directions to the output device. The mechanical gear arrangements of document US 2004251061 (A1), namely the differential actuators imply an inherent risk of jamming.
Electro-mechanical actuators comprise motors usually rotating at high speed. It needs a gearbox to transmit the high rotation speed into high torque. The usage of gearboxes brings important issues concerning the reliability. In automotive industry, such systems are used for servo steering systems. The safety requirements in automotive against mechanical jam are lower than for aerospace, which explains that they use motors with gearbox without problems. To apply such architecture to aerospace, the only solution would be to bring mechanical redundancies within the actuators and thus increase weight and complexity. Therefore this state of the art in automotive industry is not suitable for a helicopter.
The companies Agusta/Westland disclose (http://www.agustawestland.com/node/3307) a fly by wire system that utilises electric actuators in place of hydraulic units, to provide the control inputs to the helicopter's rotor systems. Unlike any other fly-by-wire system developed for helicopters the Agusta/Westland system uses electro-mechanical actuation technology for all flying controls, i.e. the main and tail rotors. The brushless electric motor actuators incorporate quadruple four lane architecture with fail technology that allow the system to function safely even after failure of two of the systems. The replacement of simplex mechanical flying control systems with quadruple electronic and mechanical systems, replaces complex hydraulic systems by complex electronic and mechanical systems. In the event of malfunction with the pilot command usually 3 or 4 sensors are needed for position and 2 computers for analysing the signals with all needed reliability adding to costs and complexity of such a control system.
The document U.S. Pat. No. 3,735,228 discloses an electric connecting link provided in a control system with FBW architecture without electronic amplification in aircraft control. Redundancy in the control system is provided by one or more actuator channels and transmission of electric power signals eliminates the need for electronic amplification. A multiplex actuator for positioning a control element in response to a mechanical input motion includes an input transducer responsive to the mechanical motion for generating a plurality of control voltages. These control voltages are transmitted to a plurality of servomotors with each motor responsive to a separate generated voltage. Velocity couplers tied to the outputs of the servomotors produce a single rotary output equal to the velocity sum of the plurality of servomotors. This single rotary output may be used directly or converted into a linear output motion that varies in accordance with the input motion. A feedback transducer responds to the output motion and generates a plurality of feedback voltages which are interconnected with the control voltages in a balancing network for control of the individual servomotors. Particularly the multiplex actuator for producing a position output in response to an input motion includes a movable core input transformer generating a plurality of control voltages. These control voltages are transmitted to a plurality of two-phase electric motors having a fixed phase winding and a control phase winding, the latter responsive to the control voltage. Velocity couplers tied to the output of the individual channel motors produce a single rotary output equal to the velocity sum of the two-phase motors. This rotary output may be used directly or converted into a linear output motion that varies in accordance with the mechanical input motion. A movable core feedback transformer responds to the output motion and generates a plurality of feedback voltages which are connected individually to one of the two-phase electric motors in a balancing network with the input transducer. The pluralities of servomotors with separately generated voltages add to costs and complexity of the system of document U.S. Pat. No. 3,735,228. Document U.S. Pat. No. 3,735,228 does neither describe any mechanical input to a sensor nor any mechanical input to an actuator allowing feedback for close loop architecture and incorporates fly by wire architecture internally.
The document EP 1036734 A2 discloses a servo actuator apparatus having: an actuator (13) which relatively displaces an operation unit (15) on the basis of an input signal E1; a position sensor (14) which detects a relative position of the operation unit (15) and outputs a detection signal DI; an actuator (23) which relatively displaces an operation unit (25) on the basis of an input signal E2; a position sensor (24) which detects a relative position of the operation unit (25) and outputs a detection signal D2; a difference calculation unit (12) which subtracts the detection signal D2 from a command signal CI supplied from a flight control computer (11), to output the input signal EI; and a difference calculation unit (22) which subtracts the detection signal D1 from a command signal C2 supplied from a flight control computer (21), to output the input signal E2, where a positive displacement direction of the operation unit (15) is reverse and series to that of the operation unit (25), and the body units of the actuators (13, 23) are integrally movable, and the actuators (13, 23) can always operate while mutually monitoring the actual operation, and have a function that, when one of the actuators fails and falls in hardover, the other actuator immediately corrects the hardover. In this way, prevention of a hardover or suppression of the degree of a hardover can be realized by a simple configuration.
The document EP 1037130 A2 discloses a flight control apparatus for a helicopter being configured by: a control unit (10) controlled by the pilot; a steering mechanism (14) for generating an aerodynamic control force; a link mechanism (12) for mechanically transmitting an amount of control Ma in the control unit (10) to the steering mechanism (14), thereby driving the steering mechanism (14); an amount-of-control sensor (30) for detecting the amount of control Ma in the control unit (10) and for supplying a control signal Sa; a flight control law calculation unit (32) for calculating a flight control law of the helicopter based on the control signal Sa, and for supplying a driving signal Sb for the steering mechanism (14); a difference calculation unit (33) for subtracting the control signal Sa from the driving signal Sb, and for supplying a difference signal Sc; a precision servo actuator unit (20) for adding the amount of control Ma transmitted via the link mechanism (12), to an amount of difference Mc corresponding to the difference signal Sc; and the like. According to this configuration, the apparatus can be easily applied to a mechanical control transmission mechanism of an existing helicopter, and the characteristics and performance of the control transmission mechanism can be remarkably increased.
The document U.S. Pat. No. 4,112,824 A discloses a hydraulic servosystem controlling an aircraft elevator 1 including a main channel 2 which is normally operative and a stand-by channel 3 which is automatically brought into operation if the main control slide 24 of the channel 2 becomes jammed. Attempts to move the jammed slide displace an inner slide 26 against a spring 27 to connect the supply for the main channel to drain through a passageway 50, 49 and 51. Pressure in a conduit 37 thus falls and a piston 48 is displaced by the pressure in a conduit 40 connected with the stand-by channel. Valves 43, 46 operated by the piston through a linkage 41, 42 cause the stand-by channel to take over from the main channel. A linkage 12, 13, 14 actuates the control slides through pins 15 which engage members (20) FIG. 2 (not shown) pivotally connected with rods (16) on the slides, the rods having slots (15′) to accommodate the arcuate movement of the pins.
Fly by Wire is highly expensive and needs several redundancies to achieve the reliability need in aeronautics, i. e. several computers, several sensors, several communication buses for command and monitoring etc. Moreover, when the electric power shuts down, there will be no backup by Fly by Wire.